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A fuselage is the main body of a conventional aircraft, accommodating payload or equipment and connecting major assemblies such as the aircraft wings. It is part of the load-carrying structure as well as the aerodynamic shape. Not every aircraft has a separate conventional fuselage, and not every fuselage is pressurized.^[ref]

The structural role is more than enclosing a volume. Forces from attached assemblies and distributed gravitational weight must be transmitted through the body, producing bending, shear and axial loading. A pressurized body also carries loads from the pressure difference between its interior and the atmosphere. These are different loading mechanisms, even if they act on the same shell.

An original ideal model illustrates the pressure contribution. Approximate a closed cylindrical section as a thin-walled pressure vessel of radius $R$, wall thickness $t$ and internal excess pressure $\Delta p$. Force balance across a longitudinal cut of length $\ell$ gives $2R\ell \Delta p=2t\ell {\sigma }_{\theta }$. Balance of an end cap gives $\pi R^2\Delta p=2\pi Rt{\sigma }_z$. ThusThe first is hoop stress and the second axial stress. For $R=1.5\mathrm{m}$, $t=2\mathrm{mm}$ and $\Delta p=50\mathrm{kPa}$, they are $37.5$ and $18.75\mathrm{MPa}$. This stipulated membrane model excludes cutouts, joints, stiffeners and aerodynamic bending; it is not a complete fuselage stress analysis.

Shape also affects the combined flow around the body and wings. Richard T. Whitcomb's area rule work examined the longitudinal area distribution of the entire combination, rather than judging the wing and fuselage as isolated streamlined parts. That connection explains why structural packaging and aerodynamic design cannot always be optimized independently.

Richard T. Whitcomb was an aerospace engineer at the National Advisory Committee for Aeronautics and National Aeronautics and Space Administration's Langley laboratory. His research linked aerodynamic geometry with measurements of pressure, drag and flow changes near the speed of sound. NASA's official profile identifies his area rule, supercritical airfoil and winglet work.^[ref]

His 1947 sweep investigation provides a concrete earlier example. A wing was tested with zero, forward and backward sweep in the Langley 8-foot high-speed wind tunnel. Pressure measurements and wake surveys supplied loading and drag information, showing that changes associated with shocks were delayed by sweep without following a universal cosine shift exactly.^[ref]

For the area rule, the scientific question was whether the streamwise area distribution of the complete wing-body combination mattered more than judging each component separately. NASA describes testing his proposed distribution with a specially built wing-body model, then applying the result to a fighter configuration.^[ref]

An original bookkeeping example illustrates the geometric idea without predicting wave drag. If a body contributes $2.0{\mathrm{m}}^2$ of cross-sectional area at one station and the added wing contributes another $0.5{\mathrm{m}}^2$ to the stipulated combined section, their sum is $2.5{\mathrm{m}}^2$. Reducing the body's contribution to $1.5{\mathrm{m}}^2$ restores a $2.0{\mathrm{m}}^2$ total there. The relevant research concerns how that total varies along the vehicle; this one-station sum is not a drag law.

His later supercritical airfoil and winglet contributions addressed different flow mechanisms. NASA describes a shaped high-subsonic section and carefully tailored wing-tip devices, rather than treating every drag reduction as the same geometric correction.^[ref] The common experimental method is to specify the geometry, identify the predicted flow change and measure the resulting aerodynamic response.

Wave drag is the component of aerodynamic pressure drag associated with the compressible wave field produced by a moving body, including the effects of shock waves. It becomes important in transonic flow and supersonic flow. A freestream Mach number below one does not exclude locally supersonic regions and shocks around an accelerating flow over an airfoil.^[ref]

Wave drag can occur at zero lift, because a body of finite thickness displaces the flow. Its identification therefore differs from that of lift-induced drag, which depends on the finite wing's loading and wake. A measured high-speed increase in total drag can also include changed flow separation and other losses; attributing the entire increment to one mechanism requires an explicit decomposition.

Richard T. Whitcomb's report compared wing/body combinations with bodies of revolution having the same axial area distribution. Compensating for the wing's area by indenting the body substantially reduced the zero-lift drag-rise increment near sonic speed.^[ref] The related NASA account describes testing the proposed area rule using a specially prepared wing/body model in the Langley transonic tunnel.^[ref] The observations support a design principle for the tested regime rather than proving that a shape's area distribution determines all of its drag.

For a stated decomposition, the wave-drag coefficient is $C_{D,w}=D_w/(qS)$, using common dynamic pressure $q$ and aerodynamic reference area $S$. An original illustrative estimate with $q=20,000$ Pa, $S=0.30$ m${}^2$ and stipulated $C_{D,w}=0.012$ givesAt a speed of 250 m/s, overcoming that component alone requires mechanical power $D_{w}U=18$ kW. These are constructed values, not measurements from Whitcomb's report.

A useful apparatus comparison retains force-balance data, surface pressure measurements and flow visualization while recording Mach number, Reynolds number and lift. Observing a shock locates part of the wave field; it does not, by itself, measure the integrated wave-drag force.